Rotary wing aircraft with undercarriage structure

ABSTRACT

A rotary wing aircraft ( 1 ) such as a helicopter comprises a force-transmission structure ( 5 ) for transmitting forces, in particular those coming from an undercarriage ( 2 ), and that, in longitudinal elevation view, presents an L-shape in which the arrangement ( 6, 19 - 21 ) forms a transversely centered box ( 6 ), such that the structure ( 5 ) transfers the forces, at least in part, to top and bottom working covering segments of the fuselage.

The invention relates to a force-transmission structure for a rotarywing aircraft fuselage such as a rotorcraft. The invention also relatesto such a structure provided with a landing gear, and it also relates tosuch an aircraft.

BACKGROUND OF THE INVENTION

In numerous such aircraft, in particular for the purpose of ensuringthat the aircraft can be used for multiple purposes in spite of theconstraints set out below, the following are guaranteed:

-   -   ground clearance should be reduced as much as possible;    -   safety should be acceptable in the event of an emergency or        crash landing;    -   aerodynamic resistance should be low; and    -   the aircraft should have excellent stability on the ground.

These constraints lead in particular to providing landing gear mountedon the aircraft so as to be retractable, with the landing gear alsooperating in a specific manner in the event of an emergency, and inparticular of a crash.

Furthermore, in certain aircraft, it is necessary for the landing gearto be compatible with naval requirements, i.e. for landing on a ship. Insuch circumstances, the landing gear must be steerable through 180° ineither direction (±180°), it must be capable of being locked relative toits steering axis, and it must avoid a parasitic wobbling phenomenonknown as “shimmy”.

In the field of rotary wing aircraft, the term “shimmy” designates rapidoscillation of a wheel about a pivot axis leading to strong interferingforces that can even lead to the undercarriage being destroyed.

In addition, the arrangement of a nose undercarriage must not impede theoperation and the installation on the aircraft of detection equipmentsuch as radars.

The fuselage of such an aircraft, or at least some of its structuralelements, must also be compatible with the above-mentioned constraints.

Firstly, this is often incompatible with the occupants of the aircrafthaving good visibility between the passenger compartment and thecockpit.

Secondly, it is useful for the inside of the cabin to present anarrangement that allows its occupants to move without difficulty betweenthe passenger compartment and the cockpit.

It is also desirable for the cabin to present an arrangement on itsinside that leaves space available either for elements forming part ofthe aircraft, or for objects that are needed by the crew or thepassengers.

However, in the past, the structures of fuselages and undercarriageshave not enabled the desired results to be achieved.

At least, at present, complying with certain specifications iscontradictory to achieving other functions, thus making it necessary tofind compromises that are not always acceptable in practice.

With reference to the structures of rotary wing aircraft, mention ismade below of a few documents of interest.

European patent No. 1 052 169 describes, in a helicopter, a truss ofbeams rigidly secured to the floor and to the skin at the bottom of thefuselage.

European patent No. 1 426 289 describes a helicopter structure thatwithstands impacts, and an energy absorber.

That structure forms a frame constituted by tubes of fiber-reinforcedcomposite material, instead of I-section beams. The tubes are filledwith foam material. The tubes are placed under the side wall at pointswhere loads are induced in the event of a crash.

That document does not provide for a box having solid walls to beinterposed at a distance from the side walls of the fuselage, above theauxiliary nose landing gear which itself disposed close to the center ofthe aircraft in its transverse direction.

French patent No. 2 629 045 describes a structural assembly for a lightairplane. A passenger compartment is constituted by a central or mainpartition at the back of the compartment, by a front partition at thefront end of the compartment, and by a hollow central beam extending inthe longitudinal direction of the airplane.

That assembly is reinforced by the fuselage walls. On the rearpartition, there are secured the rear portion of the fuselage and a mainlanding gear.

The front portion of the central beam is then used to receive a portionof the landing gear in the retracted position.

French patent No. 2 693 976 describes a helicopter fuselage having acentral structure and, connected thereto: a front structure; a rearstructure; and a landing gear.

Those structures support a transmission unit, a main rotor, and anengine.

The central structure has a framework provided with covering elementsthat define the outside shape of the fuselage. The framework issubstantially in the form of a hexahedron with framework panels beingassembled to one another.

British patent No. 724 999 describes a frame for an aircraft such as ahelicopter. The frame forms a truss of tubes leaving large sideopenings. In order to form the fuselage, a skin covers the truss oftubes and a floor is placed on the bottom portion of the truss.

U.S. Pat. No. 4,593,870 describes a fuselage structure for an aircraft,in particular a helicopter. Under the floor of the passenger cabin,behind the cockpit, that structure has a truss of beams made ofcomposite material. The truss makes it possible to obtain increased andprogressive resistance in the event of a crash landing.

Transverse walls with central openings separate the cockpit at the frontfrom the passenger cabin behind it.

U.S. Pat. No. 5,451,015 describes an aircraft fuel tank, specificallyfor a helicopter. The tank is reinforced so as to withstand an emergencylanding, and it is located behind a solid transverse partition.

That partition, which is placed behind seats, is provided with ribs soas to contribute to withstanding loads during normal operation of theaircraft.

International patent WO 00/05130 describes a helicopter fuselage. Thatfuselage has a central portion with a front portion and a rear portionconnected thereto.

The central portion is provided with means for transmitting forces dueto the main transmission unit, to the main rotor, to the landing gear,and to the tail boom.

That document does not describe a fuselage structure having a cavity forhousing landing equipment, nor does it describe a fuselage suitable forproviding passengers and crew members with free passage in alongitudinal direction on either side of the fuselage.

With reference to undercarriages, mention is made below of a fewdocuments of interest in this technical field.

French patent No. 2 608 242 describes a shock absorber for the landinggear of a rotary wing aircraft.

That shock absorber is for main landing gear including a rocker arm. Itis then disposed in a position that is substantially vertical, and as aresult its shock-absorbing function, i.e. resilient and dampedabsorption of the energy of the downward movement of the aircraft whilelanding and touching the ground, is performed in compression under load.

French patent No. 2 635 498 describes a device for steering the wheelsof an aircraft nose undercarriage. A rod slides along its axis in a tubewith its free end carrying the wheels.

A scissors connects the tube and the rod in pivoting. A retractionactuator includes means for limiting forces in the event of a crash. Ashock absorber presents a longitudinal axis that coincides with thewheel axis.

French patent No. 2 647 170 describes a device for reducing theflexibility of a shock absorber for a helicopter undercarriage. A shockabsorber cylinder includes a shock absorber piston and a shock absorberrod mounted to slide in leaktight manner relative to the cylinder anddefining within the shock absorber a compression chamber containing ahydraulic fluid that is substantially incompressible.

An expansion chamber contains a hydraulic fluid adjacent to acompressible fluid under pressure, and communicates with the compressionchamber via means for throttling the fluid expelled from the compressionchamber.

French patent No. 2 677 951 describes an electrical steering device foran aircraft undercarriage. An electric motor is secured to a shockabsorber box and is associated with a drive shaft which is disposedparallel to the axis of the box. A steering rod is secured to the driveshaft and to a pivoting tube.

French patent No. 2 689 088 describes a shock-absorbing actuator for ahelicopter. It includes a function of limiting forces in the event of acrash. It comprises a strut and performs the functions of maneuver,absorbing shocks, and peak-limiting forces.

French patent No. 2 684 957 describes a peak-limiting device for a shockabsorber for a helicopter landing gear. That shock absorber is hingedvia ball joints. The landing gear is neither steerable nor retractable,and the peak-limiting device is always integrated in the shock absorber.

British patent No. 527 994 describes a device for steering an aircraftundercarriage. The steering axis of the wheel intersects its axis ofrevolution.

A hinge allows a rocker arm to rock relative to the structure of theaircraft. That device comprises two arms extending in elevation oneither side of a sleeve for receiving the steering axis of the wheel, inorder to position the height of said steering axis, i.e. position it inelevation.

U.S. Pat. No. 2,493,649 describes a drive system for steerable frontwheels of an aircraft with a shock absorber that presents a longitudinalaxis intersecting the axis of the wheel.

U.S. Pat. No. 5,944,283 describes a shock absorber for an anti-crashundercarriage. Its rocker arm axis is offset.

U.S. Pat. No. 6,257,521 describes an aircraft tail wheel, in which theaxis of its rocker arm is offset in order to avoid interferingmovements.

The teaching of those documents in particular does not make it possibleto obtain a nose landing gear and/or a structure for the fuselage of arotary wing aircraft for ensuring the following features simultaneously:

-   -   low ground clearance;    -   good safety in the event of a crash;    -   minimum aerodynamic resistance and instability;    -   landing gear that is unaffected by the instability phenomenon        known as “shimmy”;    -   landing gear that is compatible with naval requirements;    -   a structure of an arrangement that does not interfere with        radars or other detectors;    -   the occupants have good visibility towards the front;    -   it is easy for occupants to move about in the aircraft; and    -   inside spaces are made available.

OBJECTS AND SUMMARY OF THE INVENTION

To this end, the invention provides a force-transmission structure for arotary wing aircraft such as a helicopter, the structure comprising atleast:

-   -   a rigid arrangement extending substantially in an elevation        direction;    -   a structural bottom floor rigidly connected to the arrangement;        and    -   a machine top floor rigidly secured to the arrangement close to        the top ends of said arrangement.

According to the invention, in longitudinal elevation view the structurepresents an L-shape in which the arrangement forms a transverselycentered box, and wherein the top and bottom covering segments are ofworking type such that the forces introduced in particular by said noseundercarriage are distributed within the structure and are transmitted,at least in part, by the structure to the top and bottom segments.

In an embodiment:

-   -   the box and at least two lateral arms of the arrangement extend        substantially in the elevation direction, while the structural        bottom floor extends from the box and the lateral arms        substantially longitudinally in a forward direction;    -   the box presents a profile that is at least partially open in        section in a longitudinal and transverse plane, with at least        one rearwardly-open opening to provide at least one housing; and    -   the lateral arms of the rigid arrangement are arranged        transversely at a distance from the box so as to leave        transversely on either side of the box a passage for movement        and visibility.

In an embodiment, this structure comprises at least:

-   -   within the structural bottom floor at least:        -   two longitudinal spars extending longitudinally forwards            from a bottom base of the arrangement; and        -   two transverse spars disposed substantially transversely at            right angles relative to the box, one of these spars being            longitudinally in front of the box and the other spar being            substantially in register with a rear face of the box;    -   and within the arrangement, two lateral arms extending        substantially in elevation upwards from the rear of the        structural bottom floor on either side of the arrangement.

In an embodiment, the box and/or the structural bottom and machine topfloors are made at least in part out of composite material, e.g. out ofelements formed by carbon/carbon sandwiches connected together by rivetsand/or by adhesive.

In an embodiment, at least one anchor point for receiving a hinge of anundercarriage oscillating system is provided within the structuralbottom floor.

In an embodiment, the structural bottom floor defines, at least in part,an undercarriage retraction well that is arranged longitudinally fromthe front transverse spar of the structural bottom floor comingsubstantially into register with a rear face of the box and/or possessessubstantially the same transverse dimensions as the box.

In an embodiment, the structure is for receiving at least one steerablenose undercarriage for a rotary wing aircraft, the undercarriage beingretractable substantially rearwards and comprising at least:

-   -   a wheel set having a wheel axle about which at least one wheel        rotates;    -   a generally longitudinally-extending oscillating system having        both a free end with the wheel axle mounted close behind the        free end, and a hinged front end that is longitudinally opposite        from the free end, the oscillating system being pivotally        mounted via at least one transverse hinge for anchoring to a        force-transmission structure for being rigidly secured to the        aircraft;        -   the wheel axle being arranged on the oscillating system via            a steering pivot, having a steering axis that substantially            intersects the wheel axle;        -   the steering pivot enabling the wheel to be steered on            either side of a position corresponding to the aircraft            running in a straight line;    -   a steering actuator connected to a portion of the steering pivot        that is constrained to pivot with the wheel axle, so as to steer        the wheel on either side of the position for running in a        straight line; and    -   a shock-absorbing actuator having a connection end hinged to the        oscillating system between the hinged front end and the free        end, the shock-absorbing actuator having a coupling end axially        opposite from its connection end, the coupling end being hinged        to the structure.

In an embodiment, in this undercarriage, the oscillating system presentsan elevation stroke about its hinged front end that is greater than ashock-absorbing stroke, and the shock-absorbing actuator possesses meansfor retracting the oscillating system beyond an abutment positionmarking the end of the shock-absorbing stroke, so that the oscillatingsystem and the shock-absorbing actuator perform both the function ofabsorbing shocks and the function of retracting the undercarriage.

In an embodiment, a transverse hinge for anchoring the shock-absorbingactuator is disposed substantially in register with the box, e.g. on thebox, or beneath it on the structural bottom floor.

In an embodiment, a connection end of the shock-absorbing actuatorconnecting it to the oscillating system is disposed substantially inregister with the box, e.g. longitudinally a little in front of a frontface thereof.

In an embodiment, the steering actuator is substantially integratedabout the steering axis, e.g. the steering actuator comprises anelectric motor having a highly geared-down drive outlet, e.g. usinggearing, with a steerably-driven bushing that is driven in steering andconnected to the portion of the steering pivot that is secured to thewheel axle, and a bushing that is stationary in steering that is securedto the oscillating system.

In an embodiment, the oscillating system is a trailed single-arm rockerthat rocks in elevation; the wheel axle of the wheel set being mountedsubstantially stationary both longitudinally and in elevation relativeto the longitudinal free end.

In an embodiment, the connection end and the coupling end of theshock-absorbing actuator are hinged via ball joints respectively to theoscillating system and to the structure in such a manner that theshock-absorbing actuator is stressed mainly along its sliding axis.

In an embodiment, the shock-absorbing actuator extends substantiallyalong an elevation direction, with its coupling end then being at thetop, for hinged connection to a portion of the structure that likewiseextends substantially in the elevation direction.

In an embodiment, the shock-absorbing actuator extends substantially inan elevation direction, possibly with a small amount of slope upwardsand rearwards, i.e. the connection end of the shock-absorbing actuatorbeing close to the free end of the oscillating system.

In an embodiment, the shock-absorbing actuator extends substantially inthe longitudinal direction, with its coupling end being for hingedconnection to a reinforced structural bottom floor of the structure,while its connection end is associated with a crank fitting or the likeof the oscillating system.

In an embodiment, the substantially longitudinal shock-absorbingactuator is designed essentially to be housed in a well of the structurefor receiving the undercarriage when retracted, the connection end ofthe shock-absorbing actuator being close to the hinged front end of theoscillating system.

In an embodiment, the shock-absorbing actuator extends substantially inthe longitudinal direction with its coupling end in front of theconnection end, said shock-absorbing actuator being designed to have itsconnection end in a well of the structure while its coupling endprojects longitudinally forwards from the well.

In an embodiment, the shock-absorbing actuator extends substantially inthe longitudinal direction and acquires a small amount of slope on beingretracted and while shock absorbing, or vice versa, either:

-   -   upwards and rearwards; or else    -   downwards and rearwards.

In an embodiment, the oscillating system is a deformable parallelogram,so that the steering pivot has its steering axis maintainedsubstantially extending in the elevation direction between high and lowpositions of the oscillating system during shock absorbing and duringretraction, the oscillating system comprising a pair of trailed armsdisposed one above the other in elevation.

In an embodiment, the wheel set possesses at least one wheel having aflange and its tread coming in elevation substantially up to the levelof the free end of the oscillating system; for example said wheel setpossesses two wheels disposed in parallel on either side of the free endof the oscillating system.

In an embodiment, the landing gear presents:

-   -   means that operate specifically in the event of a crash, and        suitable, in the event of a crash, for inhibiting the retraction        function and optionally the shock-absorption function of the        shock-absorbing actuator; and    -   force-limiter means for limiting forces in the event of a crash,        arranged so that once the retraction function and optionally the        shock-absorption function has/have been inhibited, the crash        energy causes the oscillating system to rise through the        additional amplitude and possibly also through the        shock-absorbing stroke, with forces being absorbed by the        force-limiter means.

In an embodiment, the force-limiter means are integrated at least inpart in the shock-absorbing actuator, for example being completelyintegrated therein.

In an embodiment, the force-limiter means are integrated at least inpart in the oscillating system, e.g. force-limiter means are integratedin the steering pivot, thereby providing the oscillating system with anenergy-absorbing additional stroke in the event of a crash.

In an embodiment, the means for operating specifically in the event of acrash are arranged to inhibit only the retraction of the oscillatingsystem so that energy absorption is then available over the additionalamplitude stroke, and possibly over an additional stroke provided byforce-limiter means located outside the shock-absorbing actuator.

In an embodiment, the means for operating specifically in the event of acrash are arranged to inhibit both retraction and shock absorption bythe shock-absorbing actuator so that energy absorption is then availableover the maximum stroke.

The invention also provides a rotary wing aircraft such as a helicopterincluding an undercarriage mounted on the fuselage via aforce-transmission structure.

According to the invention, an undercarriage is mounted on the structurein such a manner that the landing forces are distributed, at least inpart, within the central box between the structural bottom floor and themachine top floor.

In an embodiment, the structure is covered by a fairing of the fuselage.

For example, such a fuselage fairing comprises at least:

-   -   a working covering forming an integral portion of the fuselage        taking up the forces coming from the undercarriage; and/or    -   a transparent surface improving visibility to the outside of the        aircraft; and/or    -   a hatch for the landing gear or the like.

In particular, the working covering can correspond to top and bottomcovering segments constituting the fairing of the fuselage. In thecontext of the present invention, the fairing of the fuselage candesignate equally well the skin (or covering) of the fuselage, or else,by extension, a floor when the fairing portion that is either above orbelow the floor is open (no fairing covering respectively above or belowsaid floor). By way of example, this applies to a helicopter fuselagefor which the portion above the machine floor is constituted byremovable non-working covers.

In general manner, the use of working covering segments enables thecoverings to support normal stresses (flexion, traction, compressionstresses) or tangential stresses (shear stresses) with the mainconsequences being less stress on the remainder of the structure andconsequently an overall weight saving.

However, this situation does not exclude the possibility of causing thefloors to work in part.

In an embodiment, the equipment such as radars or detection equipment ismounted on the structure in such a manner as to ensure installation andoperation thereof on the aircraft is not impeded.

For example, a radar can be mounted to be retractable into a spaceprovided in the force-transmission structure, such as a space in thestructural bottom floor.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is described below with reference to embodiments given byway of non-limiting example and illustrated in the accompanyingdrawings, in which:

FIG. 1 is a fragmentary diagrammatic longitudinal elevation view showingan embodiment in accordance with the invention of a nose undercarriageand of a force-transmission structure, the nose undercarriage beingdrawn in dashed lines in its retraction position, and in uninterruptedlines in its extended position.

FIG. 2 is a fragmentary diagrammatic perspective view of the noseshowing a force-transmission structure together with a noseundercarriage in accordance with the invention.

FIG. 3 is a fragmentary diagrammatic elevation view in longitudinalsection showing an example of a force-transmission structure inaccordance with the invention, and the distribution of stresses withinsaid structure.

FIG. 4 is a fragmentary diagrammatic view in longitudinal elevation of arotary wing aircraft in accordance with the invention, showing anexample of a tactical radar and an infrared detector being disposedahead of the nose undercarriage.

FIG. 5 is a fragmentary diagrammatic section view in longitudinalelevation showing an example of a shock-absorbing actuator that extendssubstantially longitudinally, being hinged to a reinforced bottom floorof the structure, and to the oscillating system via a crank fitting, itscoupling end projecting longitudinally from the wheel well in a forwarddirection.

FIG. 6 is a fragmentary diagrammatic section view in longitudinalelevation showing an example of a longitudinal shock-absorbing actuatorhoused in the wheel well, the hinge of the shock-absorbing actuatorbeing close to the hinged front end of the oscillating system.

FIG. 7 is a fragmentary diagrammatic section view in longitudinalelevation showing an example of an oscillating system comprising adeformable parallelogram with trailed arms disposed one above the otherin elevation, with crash force limitation means being integrated in theoscillating system.

MORE DETAILED DESCRIPTION

In the various figures, elements that are similar are designated by thesame reference numerals.

The figures show three mutually orthogonal directions, X, Y, and Z.

X designates a longitudinal direction corresponding to the main lengthsor dimensions of the structures described.

Y corresponds to a transverse direction, corresponding to the lateraldimensions or widths of the structures described; these longitudinal andtransverse directions X and Y are sometimes said to be the horizontaldirections, for simplification purposes.

Z designates a third or “elevation” direction corresponding to theheight dimensions of the structures described: the terms high/low arerelative thereto; for simplification purposes, this Z direction issometimes said to be vertical.

Together the directions X and Y define an X, Y plane referred to as a“main” plane (perpendicular to the plane of the sheet of FIG. 1) withinwhich there is inscribed the polygon of support and a landing plane.

In the figures, reference 1 is a general reference to a rotary wingaircraft such as a rotorcraft.

In FIGS. 1 to 4, the X direction arrow points forwards relative to theaircraft 1. When a component of the aircraft 1, such as an undercarriageor landing gear 2 is placed close to a longitudinal end of the aircraft1 close to its front end, the component is said to be a “nose”component.

By way of illustration, in the example shown, the aircraft 1 is a“medium” helicopter, i.e. a helicopter weighing about 5 to 8 (metric)tonnes.

Within the helicopter 1, and as shown in particular in FIG. 2, there canbe seen a force-transmission structure 5. Such a structure 5 providesthe fuselage 3 of the aircraft 1 with a high degree of rigidity and withsafe behavior in the event of a crash.

The structure 5 forms a rigid arrangement comprising a box 6 that iscentered transversely and that extends substantially along the elevationdirection Z.

In FIGS. 1 and 2, this box 6 possesses at least one anchor point 7 forreceiving a hinge 8 of a shock-absorbing actuator 9 of the landing gear2.

Embodiments of the invention provide for two shock-absorbing actuators 9(e.g. in parallel, either horizontally or vertically), even though theexamples shown have only one shock-absorbing actuator.

The structure 5 also has a structural bottom floor 10 constituting thecabin floor.

In practice, it should be observed that a distinction is drawn between agenuine cabin floor—i.e. surfaces that serve to support occupants andcabin equipment on-board—and an “under-floor” structure that supportsthe major fraction of the mechanical forces.

The genuine cabin floor is designed to withstand puncturing and atpresent is often constituted by a leaktight composite honeycombassembly.

Longitudinal spars 14-15 and transverse spars 16-17 of the structuralbottom floor 10 are components of such an “under-floor structure”.

As explained below, the structural bottom floor 10 in the embodiments ofFIGS. 5 and 6 supports the anchor point 7 of a hinge 8 of theshock-absorbing actuator 9.

This bottom floor 10 of the cabin is provided with at least one otheranchor point 11.

This anchor point 11 receives a hinge 12 of an oscillating system 13 forsuspending and retracting the landing gear 2.

In FIG. 2, the structural bottom floor 10 possesses two longitudinalspars 14 and 15. These spars extend substantially in the longitudinaldirection X, forwards from the bottom base of the rigid arrangement orbox 6, being disposed transversely on either side of the box 6.

This structural bottom floor 10 also possesses two transverse spars 16and 17.

The transverse spars 16 and 17 are disposed substantiallyperpendicularly to the rigid arrangement and longitudinally they aredisposed:

-   -   one of them, 16, in front of the box 6; and    -   the other one of them, 17, substantially in register with a rear        face 18 of the rigid arrangement.

The transverse spar 17 visible in FIG. 1 is in register with the rearface 18 of the box 6.

In FIG. 2, it can be seen that in order to form a kind of opentransverse frame, there are provided two lateral arms 19 and 20 thatform parts of the rigid arrangement.

These lateral arms 19 and 20 of the structure 5 extend generally inelevation, i.e. substantially in the elevation direction Z, from therear transverse spar 17 in an upward direction.

In FIGS. 1 and 2, the rigid arrangement constitutes a single piece.

For this purpose, the lateral arms 19 and 20 are integrated with therear transverse spar 17 of the structural bottom floor 10, and also witha transverse top beam 21.

The top beam 21 is substantially parallel to the spar 17 but is oppositetherefrom in elevation, extending over the top ends of the lateral arms19 and 20 which it unites.

In addition, the structure 5 includes a machine top floor 22.

It should be observed that the term “floor” is used in its meaning thatis typical for rotorcraft, which for a machine top floor 22 designates aplate for supporting the main mechanical units, including, for example:

-   -   a turbine;    -   a transmission unit; and    -   a main rotor (40 in FIG. 4).

In FIG. 2, the machine top floor 22 is rigidly secured to the centralbox 6, close to the top ends of the box, the lateral arms 19 and 20, andthe top beam 21.

As can be seen in FIG. 1, when the structure 5 is seen in longitudinalelevation view (i.e. in a plane coinciding with the plane of the sheetin FIGS. 1, 3, and 4) it presents a shape that is generally L-shaped.

In this L-shape, the central box 6 forms the vertical limb of theL-shape, while the structural bottom floor 10 forms the limb at thebottom of the letter.

In other words, the box 6 and the lateral arms 19 and 20 extendsubstantially in the elevation direction Z while the structural bottomfloor 10 extends substantially from said box 6 and the arms 19 and 20substantially longitudinally (X direction) towards the front.

It is emphasized at this point that the structural bottom floor 10 isrigidly connected to the box 6, giving the L-shaped structure 5 of theinvention its ability to transmit to the airframe of the aircraft 1 theloads that come from the landing gear 2.

In structures 5 of the invention, the box 6 and/or the structural bottomand machine top floors 10 and 22 are made at least in part out ofcomposite material. In some examples, these are elements made ofcarbon/carbon sandwiches connected together by rivets and adhesive.

In FIG. 2, the box 6 presents a profile that is open at least in partboth in longitudinal section and in cross-section, having at least oneopening 23 that opens out rearwards in such a manner as to provide atleast one housing 24 (also visible in FIG. 1).

This housing 24 forms a “cupboard” of the central structure (see FIG. 4for example), which in this example is located a little behind the crewand connects the structural bottom floor 10 to the machine top floor 22.

As mentioned, this housing 24 is open via its rear face 18,longitudinally relative to the passenger cabin, so as to houseelectrical harnesses and/or avionics equipment, for example.

It should be observed that the panel of the rear face 18 is notsubjected to high levels of stress, i.e. it is not a “working” member,and as a result the presence of the opening 23 does not lead tomechanical weakness in the box 6, and thus the force-transmissionstructure 5.

At this point, it should be observed that the lateral arms 19 and 20 ofthe empty rigid frame of the structure 5 extend in the Y directiontransversely on either side of the box 6, being spaced apart from eachother.

This leaves respective internal passages 25 and 26 open on either sideof the aircraft 1 for moving and looking between the front and rearportions of the cabin, which portions are separated by the frame (17,19, 20, 21) of the structure 5.

This thus reconciles the requirement for the fuselage to be rigid, i.e.to provide safety in the event of a crash, and the requirement that waspreviously believed to be contradictory for good visibility and theability to move between portions of the cabin.

In the embodiment of FIG. 1 or 2, the force-transmission structure 5includes a well 27 substantially level with the structural bottom floor10, which well 27 is suitable for having the landing gear 2 retractedinto it, and is sometimes also known as a wheel bay.

In this example, the well 27 extends longitudinally from the fronttransverse spar 16 of the structural bottom floor 10 to level with therear end of the central box 6.

In this example, the well 27 has substantially the same transversedimensions (in the Y direction) as the box 6.

In FIGS. 1 and 2, it can be seen that the well 27 is for the most partintegrated in the structural bottom floor 10.

In particular embodiments, the well 27 is completely integrated in thebottom floor 10.

Thus, landing forces are distributed between the structural bottom floor10 and the machine top floor 22, as can be seen more clearly in FIG. 3(described in greater detail below). It should be noted that thisdistribution is also transferred, at least in part, to the top andbottom covering segments 35 and 36.

In FIG. 4, there can be seen navigation and/or tactical equipment 28,mounted on the structure 5 of the aircraft 1.

For example, this equipment 28 comprises a radar 29 retractable into aspace 30 formed within the structural bottom floor 10 in a manner thatis similar to the space for the undercarriage 2, but longitudinallynearer to the nose of the aircraft 1.

With reference to FIG. 3, there can be seen the forces and stresses thatare applied to the force-transmission structure 5, both in normaloperation of the aircraft 1 in which the structure 5 is installed, andin the event of an emergency landing.

In particular because of the way the box 6 is integrated to act as acentral structural post for the structure 5 (a back-to-front L-shape inFIGS. 1 to 3), it is possible to embed the beams of the wheel well 27 byfinding suitable bearing points against the machine top floor 22.

As a result, the structure 5 is provided with increased rigiditylongitudinally (X) towards the front of the box 6.

In the example of FIG. 4, this increased rigidity is obtained under thepilot's seat or cockpit.

Consequently, the masses that the structure 5 supports cantilevered outfrom the body 3 are also better integrated because of this improvedrigidity.

This also makes it possible to reduce the load on the two load-carryinglongitudinal spars 14 and 15.

By way of illustration, there can be seen in FIG. 3:

-   -   a moment created by a force 32 generated by the landing gear 2;    -   bottom and top shear forces 33 and 34; and    -   top and bottom segments 35 and 36 of the skin or covering of the        fuselage 3.

In this example, the force 32 generated by the landing gear 2 isdirected substantially upwards in the elevation direction Z.

The bottom shear force 33 is directed substantially rearwards along thelongitudinal direction X.

The top shear force 34 is directly substantially forwards along thelongitudinal direction X.

It can clearly be seen in FIG. 3, that by means of the L-shapedstructure 5 connected to the machine floor 22, the moment delivered bythe force 32 is balanced in part or even completely by the shear forces33 and 34 in the top and bottom covering segments 35 and 36. Naturally,the covering segments can withstand normal stresses and shear stressesunder this type of stress.

Other features of the force-transmission structure 5 are describedbelow.

In FIG. 2, it can clearly be seen that the box 6 does not come adjacentto the side walls of the fuselage 3, but leaves empty spaces via the twopassages or openings 25 and 26.

Consequently, the point where forces are applied is not close to thewalls of the fuselage 3, but is above the landing gear 2, which in thisexample is placed in a central position, being centered transversely inthe Y direction.

In FIG. 2, outside the box 6 and on a front face 37 there is installed ahinge plate 38 for the shock-absorbing actuator 9 of the landing gear 2.

Also, one or more rods 39 (only one rod 39 is shown in FIG. 2) forcontrolling the main rotor 40 (FIG. 4) are installed outside the box 6,in front of and close to its front face 37.

In FIG. 1, the plate 38 and the rods 39 are covered by a fairing 41(diagrammatically represented by a chain-dotted line).

In the embodiment of FIG. 2, the width, i.e. the dimension in thetransverse direction Y, of the box 6 or central post is substantiallyequal to that of the wheel well 27 (e.g. being about 500 millimeters(mm)), and also equal to the width of an equipment panel situatedbetween the pilot and the co-pilot.

Likewise, the transverse dimension of the box 6 is substantially equalto that of a central panel 42 of the machine floor 22 which also carriesequipment and auxiliary controls (e.g.: a fuel shut-off valve, a rotorbrake, or engine controls).

As an image, it can be said that the structure 5 is L-shapedlongitudinally while being substantially upside-down T-shapedtransversely (e.g. when seen from in front).

In an embodiment, the main parts of the structure 5 can be considered asbeing beams made up of two flanks made from rectangles.

The bottom rectangle defines the well and the vertical rectangle definesaccess to the cockpit.

Such a structure enables forces to be transmitted from the landing gear2 and enables those forces that are not taken up by the structure 5 tobe shared with the coverings of the fuselage 3.

In certain embodiments, such a central box 6 thus serves:

-   -   to hold the front of the machine top floor 22, in particular        when anti-crash troop seats are secured to said machine top        floor; and    -   stiffen the structure 5, and thus the aircraft 1, while taking        load from the load-carrying longitudinal spars 14 and 15 on        either side.

Certain practical aspects of the structure 5 should be considered interms of how the cabin is used:

-   -   since the central box 6 is relatively narrow (e.g. about 0.5        meters (m) wide) compared with the fuselage 3 in the location of        the cabin (e.g. 2 m wide), the box constitutes only a small mask        in the forward direction for the passengers, whereas a        conventional configuration with a central passage and with        equipment bays on either side obstructs the forward view, often        in unacceptable manner; and    -   via its sides, the central box 6 provides easy and fast (i.e.        practical) access to a first row of passenger seats (e.g. the        passenger seat 43 drawn in dashed lines in FIG. 4) via doors        (from inside the aircraft 1 via the passages 25 and 26) situated        longitudinally (X direction) towards the front of the aircraft        1, e.g. a front door 44 as shown in FIG. 4.

In embodiments of the invention, these practical aspects of thestructure 5 make it possible to provide cabins having four rows of seatsof modest size (e.g. 1.4 m) in elevation (Z), since all of these rowsare accessible without moving longitudinally (X) in the cabin.

Conventionally, in order to be able to move in this way, it is necessaryto have at least 0.2 m of additional elevation (Z), thus losingavailable area on the floor 10 that would otherwise be useful forinstalling seats.

In aircraft 1 of the invention and intended for military missions, theL-shaped and T-shaped disposition enables a commando leader tocommunicate (via the passages 25 and 26) with the crew and specify thebest landing site on a field of operation.

Also, the way the structure 5 is disposed leaves room for machine gunsto be used via portholes, whereas in many conventional configurations,such movements are impeded.

In addition to those aspects, FIG. 2 shows how the following itemspresent the same width:

-   -   the central box 6;    -   the well or wheel bay 27; and    -   the equipment and ceiling panels.

This makes it naturally easy to install vertically-extending electricalharnesses, hydraulic pipework, or the like.

In numerous embodiments, the dimensioning of the panels constituting thestructure 5 is designed to ensure that webs are positioned at themargins of the panels.

Furthermore, the internal dimensions of the elements of the structure 5enable avionics equipment to be housed therein, and makes it easy forthe equipment to be ventilated (e.g. by taking in air from below andexhausting it upwards).

It is also easy to ensure sealing relative to the cabin.

In embodiments of aircraft 1 in accordance with the invention that areintended for naval missions, and as shown in FIG. 4, the navigationand/or tactical equipment 28 includes an infrared detector device 45located longitudinally in front of the radar 29.

Like the tactical radar 29, the infrared detector device 45 is installedin front of the landing gear 2.

Thus, the infrared detector device 45 has a good field of view which mayeven be total in a forward direction and in a downward direction, andthe radar 29 can look over 360° in a horizontal plane in most flightconfigurations.

For this purpose, care is taken in an aircraft 1 of the inventionprovided with a radar 29 to ensure that the radar has good visibility inparticular sideways while turning and possibly also upwards a little inthe forward direction so as to enable it easily to take on weather radarfunctions.

In FIG. 4, the radar 29, which is retractable into the space 30,improves the ground clearance of the aircraft 1 and reduces aerodynamicdrag when the radar 29 is not active.

It is in particular the way in which the landing gear 2 of the inventionis compact in the longitudinal direction (X) that makes it easy toprovide a space 30 of dimensions that are sufficient to enable the majorpart of the radar 29 antenna and its radome to be retracted while flyingwith the radar off, or when on the ground.

In FIG. 4, the space 30 is disposed longitudinally between the infrareddetector device 45 and the undercarriage 2.

This is advantageous, in particular when turning, since the device 45 isthen in a low position, and because the device 45 is installed in a zoneof the fuselage 3 that presents smaller width transversely.

In addition, on the ground, neither the radar 29 nor the infrared device45 interfere with handling the aircraft 1 with a tow bar connected toits nose gear 2.

There follows a description in greater detail of the undercarriage orlanding gear 2 of the invention.

The undercarriage 2 is for use with an aircraft 1 and is retractablesubstantially by moving rearwards.

In general, the undercarriage 2 comprises at least one wheel set havinga wheel axle 47 about which at least one wheel 4 revolves with a flangeand a tread 49 formed in FIG. 1 or 2 by means of a pneumatic tire.

The undercarriage 2 also possesses an oscillating system 13 with thewheel axle 47 being mounted close to a longitudinally rear free end 51of the oscillating system 13.

Opposite from the free end 51 in the longitudinal direction X, theoscillating system 13 has a hinged front end 52 pivotally mounted on thestructure 5 of the aircraft 1.

At this point, it is recalled that the structure 5 is rigidly secured tothe aircraft 1.

In FIGS. 1 and 2, the hinge 12 mounts the oscillating system 13 to thestructure 5 to pivot about a transverse axis, and more preciselyrelative to the structural bottom floor 10, as described above.

In order to reduce or eliminate any offsets that make it difficult tosteer the undercarriage 2, provision is made for the wheel axle 47 ofthe wheel assembly 46 to be mounted on the oscillating system 13 via asteering pivot 54 (represented diagrammatically by shading in FIG. 1)with a steering axis 53 intersecting the axis of the wheel axle 47.

This steering pivot 54 makes it possible to steer the wheel 4 or allowit to swivel on either side of a position corresponding to the aircraft1 running in a straight line.

A steering actuator 55 (also represented diagrammatically by shading inFIG. 1) also forms part of the undercarriage 2.

The steering actuator 55 is connected to a portion of the steering pivot54 that is constrained to turn with the wheel axle 47 so as to steer thewheel 4 on either side of a position for running in a straight line.

In FIG. 1, the steering pivot 54 and the steering actuator 55 are bothsubstantially integrated around the steering axis 53.

In this example, the steering actuator 55 includes an electric motorwhose outlet is geared down to a great extent, e.g. with gearing.

A moving pin of the steering actuator 55 that is driven for steeringpurposes is connected to the portion of the steering pivot 54 that issecured to the wheel axle 47.

Another pin, that does not move for steering purposes, is secured to theoscillating system 13.

In an embodiment of the invention, in order to enable movements to beperformed on the ground by means of a towing tractor, the steeringactuator 55 is declutchable.

It should be observed that in spite of its relatively high cost, usingan actuator 55 presents advantages, in particular compared with aconfiguration for the landing gear 2 in which the wheel 4 is offset.Thus, it is easy to use landing gear 2 of the invention provided with anactuator 55 to ensure that:

-   -   the wheel 4 can easily be locked on the aircraft axis (in        flight) e.g. by interrupting the (electrical or other) power        supply to the actuator once the wheel 4 has been retracted into        the wheel bay or well 27 and turned to lie on the axis of the        aircraft 1; in contrast, for an ordinary landing gear with an        offset wheel, in order to achieve the same result it is        necessary to provide it with a specific locking mechanism;    -   it is easy to return the wheel 4 to the axis of the aircraft        using the actuator 55 while the wheel is being retracted into        its well or bay 27; in contrast, with ordinary landing gear        having a trailed wheel, it is necessary to add a specific        mechanism (e.g. springs) to achieve the same result;    -   the shimmy phenomenon is prevented because of the angular        stiffness specific to the actuator 55; however, ordinary landing        gear with an offset wheel is sensitive to this phenomenon; and    -   for a given aircraft 1, the longitudinal dimension of the entire        landing gear 2 of the invention can easily be made smaller than        that of ordinary landing gear having an offset wheel.

In FIG. 6, in order to avoid interference between the cylinder of theshock-absorbing actuator 9, it can be seen that the actuator 55 isinstalled further down on the steering pivot 54 (i.e. closer to thewheel axis 47) than in the embodiment of FIG. 1.

In this example, the actuator 55 is offset transversely to one side ofthe steering pivot 54, for the same purpose.

Returning to the examples of FIGS. 1 and 2, it can be seen that theshock-absorbing actuator 9 is an oleostrut.

A connection end 56 of the shock-absorbing actuator 9 is hinged to theoscillating system 13 close to the longitudinally rear free end 51between said free end 51 and the hinged front end 52 of the oscillatingsystem 13.

Conventionally, opposite from its connection end 56, the shock-absorbingactuator possesses a coupling end 57 hinged in the example of FIG. 1 tothe plate 38 that is rigidly mounted on the box 6 of theforce-transmission structure 5.

That constitutes a “rocker” arm undercarriage 2 with a trailed wheelthat is retractable and steerable.

As this stage, mention can be made briefly of various knownundercarriages and of the drawbacks they can present under certaincircumstances.

Landing gear can be straight, typically with a rod when stationary orwith a retracting actuator when retractable. In general such landinggear does not provide an anti-crash protection.

Nevertheless, in order to obtain an anti-crash undercarriage, someaircraft are sometimes provided with a device having a composite tube(e.g. made of carbon) which collapses while opposing a substantiallyconstant force. Such a device is similar to that shown in FIG. 7, asdescribed below.

With such a device, the shock absorber becomes blocked at the verticaltravel speeds typical of a crash (NB.: about 8 meters per second (m/s)to 11 m/s, as opposed to a speed that is usually less than 1 m/s duringa normal landing).

For this purpose, the undercarriage, or more generally the aircraft, isfitted with sensors and appropriate processor units which, for example,act on dedicated means 61 that are described below.

With the landing gear leg pivoting to take up a flat position, aphenomenon of the tires collapsing instantaneously can lead to largeamounts of damage to humans and/or equipment, if the landing gearstrikes the bottom floor hard.

It is appropriate to observe that retractable landing gear with a directshock absorber is suitable for helicopters of relatively high weight(e.g. in excess of 8000 kilograms (kg)), since the ground clearance ofsuch aircraft 1 is generally greater than 500 mm, and the floor of thecockpit is higher than the floor of the cabin.

However, that solution is difficult to apply to aircraft of lowertonnage, for reasons of bulk.

Thus, the minimum dimensions needed for the wheel well can be about 1450mm in length and 500 mm in height.

Nose landing gears of another type are also known. Such landing gear isoften used on low-tonnage helicopters.

A lever is then used to shorten the stroke of the shock absorberrelative to the vertical stroke of the center of the wheel 4.

Such landing gear can be non-retractable or retractable.

With this type of landing gear it is still not possible to reduce thedimensions of the wheel bay or the height of the landing gear whenextended, in order to fit in a given floor.

The floor lies under the cockpit and extends the floor of the cabin, andground clearance is proportional to the size of the aircraft.

The minimum dimensions of the wheel well are clearly smaller when notusing such landing gear, for example about 1100 mm in length and 400 mmin height.

Because of an offset between the axis of the wheel and its steeringpivot that is often large, this type of landing gear makes it easy toturn while running on the ground by acting on the tail rotor or bybraking the main wheels differentially. Such landing gear is alsocompact.

However it leads to a weight penalty and a penalty in terms ofmanufacturing costs.

Furthermore, its position set back rearwards relative to the aircraftgenerally requires overall stability to be restored at the main landinggear.

In practice, such landing gear is used only when anti-crash functionsare not required, since it leads to the same difficulties as straightlanding gear.

In addition, this kind of landing gear runs the risk of the wheelstaking up a sideways position under the effect of lateral forces andstriking one of the beams of the wheel bay making it impossible toretract the landing gear leg.

The specific features and advantages of landing gear in accordance withthe invention can be better understood in the light of the above.

In FIG. 1, it can be seen that the undercarriage 2 of the invention isarranged in such a manner that the oscillating system 13 moves through amaximum elevation stroke 58 about its hinged front end 52 that isgreater than its shock-absorbing stroke 59 (see FIG. 1 in particular).

In other words, when operating on the ground, the shock-absorbingactuator 9 allows the oscillating system 13 to move through ashock-absorbing stroke 59 between a low position and an end-of-shockabsorption or top abutment position.

Such a shock-absorbing stroke 59 is provided for the purposes ofstabilizing the aircraft 1 and also of providing comfort.

When retracting, the shock-absorbing actuator 9 allows a total elevationstroke 58 that is said to be a “maximum” stroke.

The elevation stroke 58 is to allow the undercarriage 2 to be retractedinto the well 27.

Thus, the amplitude of the elevation stroke 58 is greater than that ofthe shock absorbing stroke 59 for operating on the ground, by an amountreferred to as the additional amplitude and referenced 60 in FIG. 1.

This additional amplitude for the undercarriage 2 extends from theend-of-shock-absorbing position to a retracted position.

The high and low end-of-stroke abutments for shock absorbing andretracting are not shown. In some embodiments, a shock-absorbing stroke59 can be provided (in elevation in the Z direction) that is about 200mm.

For retraction purposes, the shock-absorbing actuator 9 possesses means61 that are dedicated thereto, that enable it to travel through a strokebeyond the position of the high abutment of the oscillating system 13.Activating these retraction means 61 cause the landing gear 2 to beretracted into the well 27.

In the event of a crash, sensors and processor units of the aircraft 1can act on the dedicated means 61 in order to inhibit retraction of theshock-absorbing actuator 9 and limit the forces transmitted to thestructure 5.

It can thus be understood that in this undercarriage 2 in accordancewith the invention, the oscillating system 13 and the shock-absorbingactuator 9 perform both the function of absorbing shocks and thefunction of retracting the undercarriage 2.

These functions can be associated with a specific anti-crash function,as described below.

In FIGS. 1 and 2, the oscillating system 13 is a trailed single-armvertical rocker arm, as compared with landing gear 2 comprising adeformable parallelogram.

In this example, the wheel axle 47 of the wheel set 46 is substantiallystationary in the longitudinal direction and in the elevation directionrelative to the longitudinally free end 51.

In addition, when the single arm is rocking functionally, the wheel axle47 follows a curvilinear shock-absorbing path.

In these examples, it should be observed that the connection end 56 andthe coupling end 57 of the shock-absorbing actuator 9 are hinged viaball joints, respectively to the oscillating system 13 and to thestructure 5.

The shock-absorbing actuator 9 in this example is subjected to stressmainly along its sliding axis 62 (FIG. 1).

In FIG. 1, the sliding axis 62 and the strut of the shock-absorbingactuator 9 extend substantially in the elevation direction Z, with itscoupling end 57, then referred to as its top end, being hinged to thecentral box 6 of the structure 5.

To be more precise, as can be seen in FIG. 1 the shock-absorbingactuator 9, which in this case extends substantially in the elevationdirection Z, slopes a little relative to said direction Z, slopingrearwards on going upwards.

In contrast, in the examples shown in FIGS. 5 and 6, the shock-absorbingactuator 9 extends substantially in the longitudinal direction X.

To summarize, the undercarriage 2 having a substantially longitudinalshock-absorbing actuator 9 in FIGS. 5 and 6 has its coupling end 57hinged to the reinforced structural bottom floor 10 of the structure 5,while the connection end 56 thereof is associated with a crank fitting64 or the like of the oscillating system 13.

In FIG. 6, the shock-absorbing actuator 9 is housed in the well 27 andthe connection end 56 is close to the hinged front end 52 of theoscillating system 13.

In contrast, in FIG. 5, the shock-absorbing actuator 9 has its couplingend 57 in front of the connection end 56, which connection end isdisposed in the well 27.

The coupling end 57 projects longitudinally from the well 27 forwardsthe X direction (i.e. to the left in FIG. 5).

This is similar to the situation shown in FIGS. 1 and 2 in which theshock-absorbing actuator 9 extends substantially in elevation (Z) andprojects upwards beyond the well 27.

It should be observed that such a shock-absorbing actuator 9 takes on asmall slope on being retracted or on absorbing shock, or vice versa,either:

-   -   upwards and rearwards; or else    -   downwards and rearwards.

When the structure 5 (i.e. the floor 10 in FIG. 1, and a front partitioncomparable to the transverse spar 16 in FIG. 5) possesses an opening forpassing the shock-absorbing actuator 9, the dimensions of the openingmust be sufficient to allow the shock-absorbing actuator to tilt in thisway.

On the same lines, it can be seen in FIG. 2 that the hinge 12 of theoscillating system 13 lies outside the well 27. The transverse spar 16has an opening for passing the front of the oscillating system 13.

When they are on top (e.g. in a “horizontal” partition like the floor10), such system-passing openings can contribute to reducing the groundclearance of the aircraft 1.

For a front system-passing opening (e.g. in a “vertical” partition suchas the transverse spar 16), it is possible to adjust the longitudinal(X) position of the undercarriage 2 as a function of the stresses towhich the aircraft 1 is subjected, such as those involved in achievingbalance on the ground.

Compared for example with a known landing gear, the invention thusenables the landing gear to be positioned better.

Also, a front system-passing opening, by disengaging the wheel well 27,can contribute to reducing the ground clearance of the aircraft 1.

In FIGS. 5 and 6, the coupling end 57 is hinged to a reinforced member65 of the force-transmission structure 5.

This member 65 is distinct from the box 6, and forms part of atransverse front face of the well 27. In this case it is integrated inthe structural bottom floor 10.

It can also be seen in FIGS. 5 and 6 that the connection end 56 isassociated with a crank fitting 64 or the like, itself pivotally mountedon the member 65 via a transverse hinge 12.

In the embodiment of FIG. 6, the substantially longitudinalshock-absorbing actuator 9 is arranged so as to be housed essentially inthe well 27 into which the undercarriage 2 is retracted.

In the embodiment of FIG. 5, the shock-absorbing actuator 9 extendssubstantially longitudinally with its coupling end 57 in front of itsconnection end 56.

In the embodiment of FIG. 6, the shock-absorbing actuator 9 extendslongitudinally with its coupling end 57 behind its connection end 56.

During retraction of the shock-absorbing actuator 9, it pivots slightly,such that during certain stages of its operation it presents thefollowing:

-   -   in the example of FIG. 6, a small slope that is upwards and        rearwards; and    -   in the example of FIG. 5, a small slope that is downwards and        rearwards.

In the variant shown in FIG. 7, the oscillating system 13 is adeformable parallelogram so that the steering pivot 54 has its steeringaxis 53 maintained to extend substantially along the elevation directionZ between the high and low abutment positions of the oscillating system13.

In FIG. 7, the deformable parallelogram comprises a pair of trailed arms66 disposed one above the other in elevation (Z).

Each arm 66 is connected to the structure 5 via an anchor hinge 12 thatis separate from the hinge of the other trailed arm.

Similarly to the examples of FIGS. 5 and 6, in FIG. 7, the hinges 12 arearranged on a transverse front face of the landing gear well 27.

It should also be observed in these three embodiments that thetransverse anchor hinge 8 of the shock-absorbing actuator 9 is disposedsubstantially in register with the box 6, and more precisely:

-   -   on the box 6 in FIG. 7;    -   under the box 6 and connected to the structural bottom floor 10,        in FIG. 5; and    -   slightly in front of the front face 37 of the box 6 (in the        longitudinal direction X) in FIG. 5.

Returning briefly to FIG. 1, it can be seen that the wheel set 46possesses at least one wheel 4.

Its flange and its tread 49 come in elevation (Z) substantially up tothe rear free end 51 of the oscillating system 13.

In the example of FIG. 2, it can be seen that this wheel set 46possesses two wheels 4 located in parallel with a substantially commonwheel axle 47, these wheels 4 being disposed transversely along the Ydirection on either side of the rear longitudinal free end 51.

Such a configuration is useful for minimizing the diameter of the wheels4 relative to the vertical dimension of the well 27.

According to an advantageous aspect, the undercarriage 2 of the figuresalso presents a specific mode of operation in the event of a crash.

To summarize, it can be understood that the undercarriage 2 presents:

-   -   means 50 that operate specifically in the event of a crash, that        are adapted in the event of a crash to inhibit the retraction        function and possibly also the shock-absorption function of the        shock-absorbing actuator 9; and    -   force-limiter means 67 for limiting forces in the event of a        crash and arranged so that once retraction and possibly also        shock absorption has been inhibited, the energy of the crash        causes the oscillating system 13 to rise through the additional        amplitude 60, or the shock-absorbing stroke 59, with forces        being absorbed by said force-limiter means 67.

In embodiments, the inhibitor means 50 and the force-limiter means 67are integrated in the shock-absorbing actuator 9.

However the force-limiter means 67 could be integrated in part in theshock-absorbing actuator 9 and in part in the oscillating system 13.

In embodiments that are not shown, provision can be made for the means50 and/or 67 to be distinct from the shock-absorbing actuator 9, e.g.being disposed in parallel therewith or at the hinges 12 between thestructure 5 and the oscillating system 13.

For example, in FIG. 7, the force-limiter means 67 and the inhibitormeans 50 are integrated in the steering pivot 54. In this example, themeans 50 and 67 provide the oscillating system 13 with an additionalstroke 68 for absorbing energy during a crash.

In this example other force-limiter means 67 and other inhibitor means50 form parts of the shock-absorbing actuator 9.

In other embodiments, the inhibitor means 50 are arranged to act in theevent of a crash to inhibit only retraction of the oscillating system13.

Energy absorption is then available over the stroke 60 of additionalamplitude and possibly also over the additional stroke 68 offered by theforce-limiter means 67 located outside the shock-absorbing actuator 9.

Nevertheless, it can be practical for the specific operation inhibitormeans 50 to be arranged in the event of a crash to inhibit bothretraction and shock absorption by the shock-absorbing actuator 9.

Under such circumstances, as in FIG. 1, energy absorption is madeavailable over the maximum stroke 58.

In other words, in various examples, the means 67 possess elementsand/or functions that are shared with the retraction means 61.

Under such circumstances, the shock-absorbing actuator 9 is in a lockedstate and its mechanical behavior is comparable to that of a rod or astrut. In the rather common event of the shock-absorbing actuator 9being hydraulic, throttles or other fluid flow obstructions enable suchbehavior to be obtained.

In other embodiments, the force-limiter means 67 are distinct from theshock-absorbing actuator 9.

Thus, in FIG. 7, an internal cylinder having the wheel axle 47 of thewheel 4 mounted thereon is held by means 67 in the form ofcontrolled-breakage studs that allow the cylinder to rise in the eventof a crash.

The magnitude of the crash stroke may be 40 mm, for example.

Other means 67, e.g. means that provide controlled plastic deformation,can be provided in embodiments of the invention that are not shown.

In the invention, it is thus possible to obtain a rotary wing aircraft 1such as the helicopter 1 that includes an undercarriage 2 mounted to itsfuselage 3 via a force-transmission structure 5.

The undercarriage 2 and the structure 5 are then arranged in such amanner that landing forces are distributed within the structure 5 by thebox 6 between the structural bottom floor 10 and the machine top floor22.

In an embodiment, the structure 5 is covered by a fairing forming partof the fuselage 3.

For example, such a fairing of the fuselage 3 comprises at least:

-   -   a working covering, forming an integral portion of the fuselage        3; and/or    -   a transparent surface providing better visibility to the outside        of the aircraft 1; and/or    -   a hatch, for the landing gear or the like.

In helicopters, having transparent fairing surfaces in the fuselage 3 isvery useful for visibility towards the ground, particularly whenlanding.

As explained above, equipment (such as radars 29 or an infrared detectordevice 45) is mounted on the structure 5, so as to avoid interferingwith installation and operation thereof on the aircraft 1.

Such an aircraft 1 is multipurpose since it can perform the followingmissions amongst others:

-   -   transporting civilian passengers, including to and from offshore        platforms;    -   search and rescue;    -   medical emergencies;    -   transporting military troops;    -   identifying and monitoring surface vessels or submarines.

As mentioned above, specifications that are favorable to all of thesemissions includes in particular:

-   -   moderate ground clearance for the aircraft 1, by obtaining a        compromise between the proximity of the bottom of its fuselage 3        with an unprepared landing ground (X, Y);    -   direct accessibility to the cabin for commandos or accessibility        via simple steps for passengers;    -   lateral stability (crosswinds, on deck);    -   anti-crash capacity in the landing gear 2 and in the fuselage 3;        and    -   undercarriage 2 that is retractable so as to reduce fuel        consumption and/or increase the speed of the aircraft 1 on long        missions (e.g. of more than 3 hours), while leaving the radar        with a field of view that is clear over 360°.

With the invention, it is also easy to provide a polygon of support (atriangle for aircraft 1 having three undercarriages, including anundercarriage 2) to be well located relative to the center of gravity ofthe aircraft 1.

This is made possible by the fact that the main wheels 4 of the aircraft1 are as close as possible to the center of gravity (from behind)without that leading to any risk of the aircraft tilting onto its tailwhen the cabin is empty.

The invention also makes it possible to ensure that the nose wheels 4are far enough forwards, while nevertheless not being too far forwardsso as to avoid deforming the cabin on the ground or while landing (forexample leading to cracks in the canopy).

FIG. 2 shows a clear example of the structure 5, the undercarriage 2,the flight controls, and the avionics equipment being integrated.

The structure 5 in this example comprises the box 6 in the form of achannel-section beam having its longitudinal axis extendingsubstantially vertically.

The shock-absorbing actuator 9 is secured to the outside of the web ofthe channel-section box 6, on the front side thereof relative to theaircraft 1.

This provides solutions when the shock-absorbing actuator 9 needs to becapable of penetrating into the cockpit and needs to have a structurallystrong attachment point therein; under such circumstances, a structure 5designed in accordance with the invention is appropriate.

When faced with problems of stability on the ground, the solutionretained herein is to steer the wheels 4 without them being offset to alarge extent, by using an electrical actuator (that can be declutched toallow towing on the ground).

This also presents the advantage of being suitable specifically for useon the deck of a ship where it is necessary to be able to turn aroundthe docking harpoon in order to become aligned with the hangar prior tostopping the engine (and the rotors).

With aircraft 1 of a certain tonnage (e.g. more than 7 tonnes), fittedwith nose gear of the straight type, it is necessary to develop priorart landing gear, i.e. it is necessary to undertake development that isspecific to naval use without any offset between the wheels and thesteering pivot associated with a steering actuator, in order to enablethe helicopter to be moved on the deck of a frigate.

Furthermore, with the invention it is possible to provide a nose gear 2that is very compact at its bottom end: the well 27 for receiving thelanding gear 2 need have a length of only about 950 mm while stillmaintaining 400 mm of ground clearance.

Behavior during a crash becomes entirely appropriate and predictablesince:

-   -   the wheels 4 are held in the straight position by the actuator        and therefore do not run the risk of entering the well 27 for        the landing gear 2 while in a crosswise position; and    -   the vertical extent of the well 27 is determined so that the        bottom of the structure 5 can be flattened during a crash        without being hindered by the wheels 4: they are free to rise        far up into the well 27.

Without major difficulty, the shock-absorbing actuator 9 can combine thefunctions of providing normal shock absorption, crash impact absorption,and landing gear retraction, particularly since the raised position ofthe wheels 4 corresponds to the position taken up by the wheels at theend of a crash.

Within an aircraft 1 of the invention, it even becomes possible to use ashock-absorbing actuator 9 of the same type (ignoring adjustments) forthe nose gear and for the main landing gear (providing it is of therocker arm type).

Finally, this type of landing gear is less expensive than a nose gear 2of conventional type, and has been found in testing to be more reliablewhen faced with piston wear and leakage.

In some embodiments, the steering actuator is an electric motor with agearbox unit located directly on the axis of the hinges.

With the invention, it is easy to provide for the landing gear 2 to besteerable through at least ±180° and for it to be retractable, and forit to be possible to lock the landing gear 2 on its axis (for straightline towing) while avoiding the shimmy phenomenon.

In variants of the invention in which the actuator 9 is installedsubstantially in a horizontal position, it is necessary to provide acrank fitting 64 close to the front attachment point of the rocker arm.

This solution avoids the actuator 9 introducing into the cabin.

However, because of the lever arm provided by the crank fitting, theforces generated are amplified and it becomes necessary to reinforce theattachment points on the structure 5 for the shock-absorbing actuatorand for the rocker arm.

1. A rotary wing aircraft comprising a crew cabin, a passengercompartment rearward of the cabin in a direction from a nose to a tailof the aircraft, a force-transmission structure below the cabin, asteerable nose undercarriage connected to said force-transmissionstructure, and a fairing of a fuselage formed, at least in part, by topand bottom covering segments that are adjacent to saidforce-transmission structure, the force-transmission structurecomprising: a rigid arrangement extending substantially in an elevationdirection; a structural bottom floor rigidly connected to a bottom ofthe rigid arrangement; and a machine top floor rigidly secured at a topof the rigid arrangement; wherein, in longitudinal elevation view, theforce-transmission structure forming an L-shape, wherein the arrangementincludes a transversely centered box, and wherein the top and bottomcovering segments are such that the forces introduced by said noseundercarriage are distributed within the force-transmission structureand are transmitted, at least in part, by the force-transmissionstructure to the top and bottom covering segments, and wherein thestructural bottom floor comprises: two longitudinal spars extendinglongitudinally forwards from a bottom base of the arrangement; and twotransverse spars disposed substantially transversely at right anglesrelative to the box, one of the transverse spars being longitudinally infront of the box and the other transverse spar being substantiallyvertically aligned with a rear face of the box; and wherein thearrangement further includes two lateral arms extending substantially inelevation upwards from the rear of the structural bottom floor on eitherside of the box, the two lateral arms are arranged at a distance fromthe box so as to form with the box a transverse passage on either sideof the box for movement between the cabin and the passenger compartmentand for visibility between the cabin and passenger compartment.
 2. Arotary wing aircraft according to claim 1, wherein: the box forming aprofile that is at least partially open in section in a longitudinal andtransverse plane, with at least one rearwardly-open opening to provideat least one housing, at least one of the box and the structural bottomfloor, and the machine top floor are made at least in part out ofcomposite material.
 3. A rotary wing aircraft according to claim 1,wherein, within the bottom floor, an anchor point is provided forreceiving a hinge of an oscillating system of the undercarriage, theundercarriage being retractable substantially rearwards and comprises atleast: a wheel set having a wheel axle about which at least one wheelrotates; said oscillating system extending generally longitudinally andhaving a first end supporting the wheel axle, and a hinged front endthat is longitudinally opposite from the first end, the oscillatingsystem being pivotally mounted via at least one transverse hinge foranchoring to the force-transmission structure in order to be rigidlysecured to the aircraft; a shock-absorbing actuator having a connectionend hinged to the oscillating system between the hinged front end andthe first end, the shock-absorbing actuator having a coupling endaxially opposite from the connection end, the coupling end being hingedto the force-transmission structure; and wherein the structural bottomfloor defines, at least in part, an undercarriage retraction well thatis arranged longitudinally from the front transverse spar of thestructural bottom floor to the rear face of the box, substantially invertical alignment with the rear face of the box.
 4. A rotary wingaircraft according to claim 3, wherein the transverse hinge foranchoring the shock-absorbing actuator is disposed substantially alignedwith a front face of the box.
 5. A rotary wing aircraft according toclaim 3, wherein the connection end of the shock-absorbing actuator isdisposed substantially longitudinally in front of a front face of thebox, and wherein the connection end and the coupling end of theshock-absorbing actuator are hinged by ball joints respectively to theoscillating system and to the force-transmission structure, in such amanner that the shock-absorbing actuator is stressed mainly along asliding axis.
 6. A rotary wing aircraft according to claim 3, whereinthe oscillating system is a trailed single-arm rocker that rocks inelevation; and wherein the shock-absorbing actuator extendssubstantially in an elevation direction, with the coupling end hinged toa top of the force-transmission structure to a portion of theforce-transmission structure that extends substantially in the elevationdirection.
 7. A rotary wing aircraft according to claim 3, wherein theshock-absorbing actuator extends substantially in an elevation directionwith a small amount of slope upwards and rearwards, the connection endof the shock-absorbing actuator being close to the first end of theoscillating system.
 8. A rotary wing aircraft according to claim 1,wherein the structural bottom floor defines, at least in part, anundercarriage retraction well that receives said steerable noseundercarriage and is arranged longitudinally from the front transversespar of the structural bottom floor, said undercarriage retraction wellpossessing substantially the same transverse dimensions as said box. 9.A rotary wing aircraft according to claim 8, wherein said undercarriageretraction well is essentially vertically aligned with a rear face ofthe box.
 10. The rotary wing aircraft according to claim 3, wherein thetransverse hinge for anchoring the shock-absorbing actuator is disposedbeneath the box on the structural bottom floor.